1. Field of the Invention
The present invention relates generally to air cooled turbine blades, and more specifically to cooling the leading edge region of the turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an aero engine used in an aircraft or an industrial gas turbine engine used in a ground based power plant, includes a turbine section with multiple rows of rotor blades and stationary vanes or guide a high temperature gas through the turbine to react with the turbine blades and drive the rotor shaft. Since the efficiency of the engine can be increased by passing a higher temperature gas into the turbine, some of the rows of blades and vanes require internal cooling air that provides convection cooling, impingement cooling and film cooling for the internal surfaces and the exterior surface of these airfoils.
The leading edge region of the airfoil is exposed to the hot gas flow in a direct head-on manner. In the prior art as shown in FIGS. 1 and 2 which shows a cut-away view of the blade, the blade leading edge is cooled by impingement cooling from the backside and film cooling through film cooling holes. The prior art showerhead film cooling includes three rows in which a middle row 11 is positioned at the airfoil stagnation point where the highest heat loads occur on the airfoil leading edge. A pressure side row 12 and a suction side row 13 form the three rows in the showerhead arrangement. The film cooling holes 11 through 13 have the same constant diameter from the inlet to the outlet of the hole. Film cooling holes for each row are arranged in an inline pattern and incline at from 20 degrees to 35 degrees relative to the blade leading edge radial surface as seen in FIGS. 3 and 4.
The disadvantage of the showerhead arrangement of the prior art FIGS. 1 through 4 design is the over-lapping of film ejection flow in a rotational environment. As a result of this type of film hole arrangement, a hot streak is shown in-between film holes as displayed in FIG. 4. Hot spots can lead to erosion of the metal substrate and therefore significantly decrease the life of the blade.
U.S. Pat. No. 6,869,268 B2 issued to Liang on Mar. 22, 2005 and entitled COMBUSTION TURBINE WITH AIRFOIL HAVING ENHANCED LEADING EDGE DIFFUSION HOLES AND RELATED METHODS disclosed a showerhead with film cooling holes having a constant diameter inlet and a diffusion outlet in which the outlet has an expansion on the inward wall, the expansion being “in a range of about 5 degrees to about 20 degrees” and the constant diameter hole being angled upward at from about 20 degrees to about 35 degrees.
It is therefore an object of the present invention to provide for a turbine blade leading edge showerhead cooling arrangement that will provide for a high leading edge film effectiveness for a high temperature blade.